1. Field of the Invention
The present invention relates to turbomachinery blades, and more specifically to a replaceable leading edge insert for a blade.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
A gas turbine engine includes two types of rotor blades, i.e. fan or compressor blades, and turbine blades. Fan or compressor blades are rotated for pressurizing air, whereas turbine blades extract energy from hot combustion gases for powering the fan or compressor. Thus, airfoils of the fan or compressor blades are differently configured aerodynamically than are turbine blades.
A fan blade is a specialized form of compressor blade which is relatively larger for initially pressurizing air, a significant portion of which is used for providing propulsion thrust for powering an aircraft in flight. Individual fan blades are suitably mounted around the circumference of a rotor disk as either discrete components, or integrally therewith in the form of a one-piece integrally bladed disk or blisk. During rotation, the blades and disk rotate and therefore generate substantial centrifugal forces which are carried by the disk. As the disk rotates during normal operation of the engine, the predominantly radial centrifugal forces are developed in the blades which cause steady tensile stress therein. In addition to steady stress, the blades are also subject to alternating stress typically caused by vibration of the blade due to the airflow thereover. Neither of these stresses is uniform, but varies over the length, width, and thickness of the blade.
A bladed rotor disk of a gas turbine engine is a very complex and expensive piece of machinery. The fan blade assembly in the aero engine is susceptible to damage from a foreign object during flight such as a bird strike. Due to an impact, a piece of the leading edge of one or more blades can be damaged and require replacement to maintain proper efficiency of the fan. In an integrally bladed rotor disk or blisk, one cannot remove the blade and replace the damaged blade with a new blade because the entire bladed disk is formed as a single piece.
In a common configuration, compressor blades include integral dovetails for being removably mounted in a corresponding dovetail slot in the perimeter of a rotor disk. This permits the individual manufacture of each blade, and the individual replacement thereof in the event of blade damage during operation. However, such bladed disks require an enlarged disk rim for limiting centrifugal reaction stresses therein around the axial or circumferential dovetail slots used for radially retaining a corresponding row of rotor blades.
A modern improvement over bladed disks in a gas turbine engine is a row of rotor airfoils integrally formed with the perimeter of a rotor disk in a one-piece or unitary blisk configuration. The blade dovetails are eliminated along with the corresponding dovetail slots in the perimeter of the disk, and centrifugal loads are carried from the individual airfoils into the corresponding disk with an inherently strong load path there between. Accordingly, blisks are mechanically stronger than bladed-disks and thusly may be more efficiently configured for reducing disk size and weight for providing additional advantages and performance of the engine.
However, since the blisk airfoils are integrally formed with the supporting disk, the airfoils are not individually removable or replaceable in the event of foreign object damage thereof. Relatively small compressor blisks have been used in commercial service for many years, and are sufficiently small that they may be simply replaced in whole in the event of excessive damage to one or more of the airfoils thereof.
Alternatively, where damage is relatively minor, the damage may be simply removed, by grinding for example, thusly leaving the airfoil with a less than original configuration. This damage removal method is unacceptable for major airfoil damage since aerodynamic performance will be severely degraded, and significant rotor imbalance therefrom may be difficult to correct with ordinary balancing procedures.
Furthermore, damage removal may adversely affect strength of the airfoil itself. A typical compressor airfoil is slender and has a crescent or airfoil profile extending axially between thin leading and trailing edges. The airfoil is cantilevered from its root, with a radially opposite tip spaced closely adjacent to a surrounding casing or shroud during operation. The airfoil is typically twisted from root to tip with a complex three dimensional (3D) configuration or contour for aerodynamically pressurizing airflow during operation.
The contoured airfoil is subject to aerodynamic and centrifugal loads during operation which result in a varying pattern of stress therein. The airfoil must thusly be designed for limiting the maximum airfoil stress for enjoying a suitable useful life during operation, and the airfoil material is typically a high strength material, such as titanium, for accommodating the substantial loads carried during operation.
In the original manufacture of the blisk, its material strength must not be reduced or compromised by the various machining processes used. Excessive temperature must be avoided which would degrade material properties. For example, the machining of the individual airfoils may be done using a milling machine or an electrochemical machine having numerically controlled multiple axes for precision movement. Material is removed from the original work piece or blank with minimal heat buildup to prevent degradation of the material strength.
Accordingly, the known repair process for compressor blisks is limited to the mere removal of airfoil damage to prevent strength reduction of the airfoil.
The advantages of using compressor blisks in a gas turbine engine are presently promoting the development of substantially larger and more expensive blisks for use in multi-stage axial compressors and low pressure fan compressors upstream therefrom. Fan blisks have relatively thick airfoils and are subject to considerably less foreign object damage than the relatively thin airfoils of compressor blisks downstream therefrom. The compressor blisks are nevertheless relatively large and quite expensive.
For example, a two stage tandem blisk includes two rows of airfoils extending from corresponding disks in a unitary assembly. Damage to any one of the blisk airfoils in either stage affects the usefulness of the entire two stage blisk. The inability to repair a two stage blisk requires the entire replacement thereof at a corresponding substantial cost.
Accordingly, it is desired to provide a method of repairing a blisk for restoring airfoils thereof to an original configuration at the repair site.
Prior art inventions such as U.S. Pat. No. 4,326,833 issued to Zelahy et al. on Apr. 27, 1982 shows gas turbine engine air cooled blade member which is separately attached to a rotor disk, where the blade can be repaired by removing from the blade at least one blade segment which comprises an edge portion, a part of the chamber wall, and a part of the platform connected with and extending from the edge portion and connected with the chamber wall. The replacement member is of substantially the same type material and of a size and shape which matches the removed blade segment. The replacement segment is bonded to the blade to form a completed blade assembly.
Another well known method of repairing damaged blades is disclosed in U.S. Pat. No. 5,755,031 issued to Baumgarten et al. on May 26, 1998 which shows a method for replacing a rotor blade on an integrally bladed rotor by removing the entire damaged rotor blade while leaving a stub portion extending out of the disk, providing a die around the stub portion, extruding a portion of the stub to form a blade platform, and attaching a new rotor blade to the platform.
U.S. Pat. No. 6,095,402 issued to Brownell et al. on Aug. 1, 2000 shows a method for the manufacture or repair of a blisk by linear friction welding, in which the method comprises providing a stub on the circumferential surface of a turbine disk and defining a weld joint surface on the radially outermost face of the stub. A new blade is applied radially to the stub and welded to the stub to form a new bladed extending from the rotor disk.
U.S. Pat. No. 6,508,000 issued to Burke et al on Jan. 21, 2003 shows a transient liquid phase bonding repair for advanced turbine blades and vanes in which a damaged portion of the leading edge of the blade is removed, and an insert is fabricated to match the removed portion. The insert is then bonded to the blade.
U.S. Pat. No. 6,568,077 issued to Hellemann et al. on May 27, 2003 discloses a blisk weld repair in which a damaged blade is repaired by removing a portion of the damaged blade by machining away from the airfoil to create a cutout or notch in the airfoil, filling the notch by welding a repair to the blade, and then machining the weld repair to restore the airfoil to substantially the original, pre-damaged configuration at the repair site.
U.S. Pat. No. 5,785,498 issued to Quinn et al on Jul. 28, 1998 and entitled COMPOSITE FAN BLADE TRAILING EDGE REINFORCEMENT shows a wide chord fan blade with a reinforced region at the blade tip that is covered by a thin metallic sheathing bonded to the trailing edge with an embodiment in FIG. 3 that uses a mechanical fastener (a threaded nut and bolt) disposed through the entire local thickness of the airfoil to further secure the metallic sheathing to the airfoil.
A disadvantage of the Zelahy et al. patent is that the leading edge replacement portion of the blade is held in place in the radial direction only by the bond strength between the replacement piece and the remaining blade. The high rotational speeds of the bladed disk produce very high centrifugal force. The bond could break and the leading edge replacement portion could separate from the blade.
A disadvantage of the Baumgarten et al. and Brownell et al. patents is that the entire damaged disk is replaced, and that the repair process cannot be performed in the field, but at a repair facility.
The Burke patent uses a replacement part that is customized to the damaged part being replaced, and thus one replacement part will not work on another blade on the same integrally bladed disk or blisk.
The Hellemann et al. invention also lacks the means to radially hold the leading edge replacement to the blade, and also must be down at a repair facility and not in the field. Also, the Hellemann repaired blade is custom fitted to the damaged blade, and therefore the leading edge replacement piece cannot be used on another blade of the same blisk.
It is an object of the present invention to provide a method of and a leading edge insert member to repair a damaged blade of a blisk used in a gas turbine engine.
It is another object of the present invention to provide a leading edge insert member that is radially fixed on the blade to prevent breaking of the bond due to centrifugal force acting on the insert from high speed rotation.
It is still another object of the present invention to provide for an insert that is standardized in shape so that one insert can be used on any blade of the blisk.
It is still another object of the present invention to allow for a damaged blade to be repaired in the field instead of a machine shop or repair facility.
Another object of the present invention is to provide for a blade insert that will allow for easy separation of parts that have been joined previously for in the field replacement.